The application described herein relates generally to gas turbine engines components, and more specifically to an apparatus for sealing the gap between adjacent turbine blade platforms.
A typical gas turbine engine has an annular axially extending flow path for conducting air sequentially through a compressor section, a combustion section, and a turbine section. The compressor section includes a plurality of rotating blades which add energy to the air. The air exits the compressor section and enters the combustion section. Fuel is mixed with the compressed air, and the resulting combustion gases mixture is ignited to add more energy to the system. The resulting products of the combustion then expand through the turbine section. The turbine section includes another plurality of rotating blades, which extract energy from the expanding air. A rotor shaft interconnecting the compressor section and turbine section transfers a portion of this extracted energy back to the compressor section. The remainder of the energy extracted may be used to power a load, for example, a fan, a generator, or a pump.
At least some known rotor assemblies include at least one row of circumferentially-spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side and a suction side connected together at leading and trailing edges. Each airfoil extends radially outward from a rotor blade platform to a tip, and also includes a dovetail that extends radially inward from a shank extending between the platform and the dovetail. The dovetail is coupled to the rotor blade within the rotor assembly to a rotor disk.
The sides of platform sections of adjacent blades in a row of blades abut each other to form a portion of the boundary defining the flow path for the air and combustion gases. Although it would be desirable to have adjacent platforms abut in a perfect sealing relationship, the necessity to accommodate thermal growth and machining tolerances results in a small gap being maintained between adjacent platforms.
In order to couple the dovetail to the rotor disk, the dovetail must be machined to be slightly smaller than the slot into which it is inserted. This causes small buffer cavities in front and behind the dovetail. During operation of the turbine, cooling air may leak from the front buffer cavity, across the top of the disk, to the buffer cavity behind the dovetail, through the gap between aft skirts of adjacent rotor blades and into the flow path of the combustion gases. Leakage of the air into the flow path of the hot combustion gases causes a loss in the engine cycle and therefore decreases the engine efficiency. It is desirable to reduce this leakage to decrease specific fuel consumption, therefore increasing engine efficiency.
Accordingly, there exists a need to provide an improved device for sealing the gap between turbine rotor blade platforms of adjacent rotating blades in a gas turbine engine.